Combustor Assembly for a Turbine Engine

ABSTRACT

A combustor assembly for a gas turbine engine includes a dome defining a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end is received within the slot of the dome. In one aspect, the forward end of the liner defines blind warming openings that allow a warming airflow to actively warm the forward end during transient operation of the engine. In another aspect, the dome defines a plurality of impingement openings that allow a warming airflow to impinge on the forward end to actively warm the liner during transient operation of the engine. In another aspect, a grommet coupling a pin with the liner defines a plurality of warming passages that allow a warming airflow to flow therethrough to warm the forward end of the liner during transient engine operation.

FIELD

The present subject matter relates generally to gas turbine engines, and more particularly to combustor assemblies for gas turbine engines.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.

More commonly, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are being used as structural components within gas turbine engines. For example, given the ability of CMC materials to withstand relatively extreme temperatures, there is particular interest in replacing components within the combustion section of the gas turbine engine with CMC materials. More particularly, inner and outer liners of gas turbine engines are more commonly being formed of CMC materials.

During normal operation, it is common for gas turbine engines to be required to rapidly increase thrust. During such rapid power increases or transient state conditions, combustion gases are generated within a combustion chamber defined by inner and outer liners. As the combustion gases flow downstream through the combustion chamber, the combustion gases scrub along the liners, causing the liners to rapidly heat up. However, the forward ends of the liners, or the portions of the liners that attach with dome sections, typically do not heat up as quickly as the rest of their respective liners. The thermal lag at the forward ends of the liners may cause undesirable bending stress and strain on the liners. As gas turbine engines may undergo many rapid power increases over many engine cycles, such repeated stress and strain on the liners can negatively impact their durability.

Accordingly, a combustor assembly of a gas turbine engine that includes features that reduce the stress and strain on combustion liners of the combustor assembly during rapid power increases of the gas turbine engine would be useful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

In one exemplary embodiment of the present disclosure, a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end of the liner defines a plurality of mounting openings spaced along the circumferential direction. The forward end of the liner also defines one or more blind warming openings.

In another exemplary aspect of the present disclosure, a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a dome comprising a yolk and a base plate spaced outward from the yolk along the radial direction, the yolk and the base plate defining a slot. Further, the combustor assembly includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot. The yolk defines a plurality of impingement openings extending therethrough such that a warming airflow impinges on the forward end of the liner during operation of the gas turbine engine.

In yet another exemplary aspect of the present disclosure, a combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction is provided. The combustor assembly includes a dome defining a slot and a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot of the dome, the forward end of the liner defining a plurality of mounting openings spaced along the circumferential direction. Further, the combustor assembly includes a plurality of mounting assemblies for coupling the forward end of the liner with the dome, each of the plurality of mounting assemblies including: a pin extending through one of the plurality of mounting openings and a grommet positioned within one of the plurality of mounting openings and defining a grommet opening, the pin received within the grommet opening, and wherein the grommet defines one or more warming passages.

These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter;

FIG. 2 provides a schematic, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure;

FIG. 3 provides a close up, cross-sectional view of an attachment point of the exemplary combustor assembly of FIG. 2, where a forward end of an outer liner is attached to an outer dome section;

FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a plurality of blind warming openings defined by the forward end in accordance with an exemplary embodiment of the present disclosure;

FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and a warming slit defined by the forward end in accordance with an exemplary embodiment of the present disclosure;

FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and impingement openings defined by the outer dome section in accordance with an exemplary embodiment of the present disclosure;

FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and depicts an impingement jacket attached to the outer dome section in accordance with an exemplary embodiment of the present disclosure;

FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and warming passages defined by a grommet in accordance with an exemplary embodiment of the present disclosure;

FIG. 9 provides a cross sectional view of the grommet of FIG. 8 taken on line 9-9 of FIG. 8; and

FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment of a combustor assembly depicting a forward end of an outer liner attached to an outer dome section and further depicting the forward end of the outer liner having enhanced surfaces in accordance with an exemplary embodiment of the present disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “forward” and “aft” refer to relative positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. It should be appreciated, that as used herein, terms of approximation, such as “about” and “approximately,” refer to being within a ten percent (10%) margin of error.

Exemplary aspects of the present disclosure are directed to a combustor assembly for gas turbine engines. In one exemplary aspect, the combustor assembly includes a dome defining a slot and a liner that at least partially defines a combustion chamber. The liner extends between an aft end and a forward end. At least a portion of the forward end is received within the slot of the dome. In one aspect, the forward end of the liner defines a plurality of blind warming openings that allow a warming airflow to actively warm the forward end during transient operation of the engine. In another aspect, the dome defines a plurality of impingement openings that allow a warming airflow to impinge on the forward end of the liner to actively warm the forward end during transient operation of the engine. In another aspect, a grommet coupling a pin with the liner defines a plurality of warming passages that allow a warming airflow to flow therethrough to warm the forward end of the liner during transient engine operation. In a further aspect, one or more surfaces of the forward end of the liner have enhanced surfaces that increase the thermal response of the forward end. By warming the forward end of the liner, the stress and strain on the liner during transient operating conditions may be reduced, particularly during transient engine power increases or bursts. Moreover, by reducing the stress and strain on the liner, the durability of the liner or liners may be improved.

FIG. 1 provides a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of FIG. 1, the gas turbine engine is a high-bypass turbofan jet engine 10, referred to herein as “turbofan engine 10.” As shown in FIG. 1, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted, the fan blades 40 extend outwardly from disk 42 generally along the radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to a suitable actuation member 44 configured to collectively vary the pitch of the fan blades 40 in unison. The fan blades 40, disk 42, and actuation member 44 are together rotatable about the longitudinal axis 12 by LP shaft 36 across a power gear box 46. The power gear box 46 includes a plurality of gears for stepping down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 is covered by rotatable spinner or front nacelle 48 aerodynamically contoured to promote an airflow through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the fan 38 and/or at least a portion of the core turbine engine 16. It should be appreciated that the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially-spaced outlet guide vanes 52. Moreover, a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 enters the turbofan 10 through an associated inlet 60 of the nacelle 50 and/or fan section 14. As the volume of air 58 passes across the fan blades 40, a first portion of the air 58 as indicated by arrows 62 is directed or routed into the bypass airflow passage 56 and a second portion of the air 58 as indicated by arrow 64 is directed or routed into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio. The pressure of the second portion of air 64 is then increased as it is routed through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 that are coupled to the outer casing 18 and HP turbine rotor blades 70 that are coupled to the HP shaft or spool 34, thus causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 that are coupled to the outer casing 18 and LP turbine rotor blades 74 that are coupled to the LP shaft or spool 36, thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the core turbine engine 16.

It should be appreciated that the exemplary turbofan engine 10 depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, the turbofan engine 10 may have any suitable configuration. For example, the present disclosure matter may be suitable for use with or in turboprops, turboshafts, turbojets, reverse-flow engines, industrial and marine gas turbine engines, and/or auxiliary power units.

FIG. 2 provides a close-up cross-sectional view of a combustor assembly 100 in accordance with an exemplary embodiment of the present disclosure. For example, the combustor assembly 100 of FIG. 2 may be positioned in the combustion section 26 of the exemplary turbofan engine 10 of FIG. 1. More particularly, FIG. 2 provides a side, cross-sectional view of the exemplary combustor assembly 100 of FIG. 2.

As shown, the combustor assembly 100 includes an inner liner 102 extending between an aft end 104 and a forward end 106 generally along the axial direction A, as well as an outer liner 108 also extending between an aft end 110 and a forward end 112 generally along the axial direction A. The inner and outer liners 102, 108 together at least partially define a combustion chamber 114 therebetween. The inner and outer liners 102, 108 are each attached to an annular dome. More particularly, the annular dome includes an inner dome section 116 attached to the forward end 106 of the inner liner 102 and an outer dome section 118 attached to the forward end 112 of the outer liner 108. The inner and outer dome sections 116, 118 may be formed integrally (or alternatively may be formed of a plurality of components attached in any suitable manner) and may each extend along the circumferential direction C to define an annular shape. As will be discussed in greater detail below with reference to FIG. 3, the inner and outer dome sections 116, 118 each include an inner surface 120 (i.e., inner relative to their respective forward ends) and a forward surface 121 at least partially defining a slot 122 for receipt of the forward end 106 of the inner liner 102, and the forward end 112 of the outer liner 108, respectively.

The combustor assembly 100 further includes a plurality of fuel air mixers 124 spaced along a circumferential direction C and positioned at least partially within the annular dome. More particularly, the plurality of fuel air mixers 124 are disposed at least partially between the outer dome section 118 and the inner dome section 116 along the radial direction R. Compressed air from the compressor section of the turbofan engine 10 flows into or through the fuel air mixers 124, where the compressed air is mixed with fuel and ignited to create the combustion gases 66 (FIG. 1) within the combustion chamber 114. The inner and outer dome sections 116, 118 are configured to assist in providing such a flow of compressed air from the compressor section into or through the fuel air mixers 124. For example, the outer dome section 118 includes an outer cowl 126 at a forward end 128 and the inner dome section 116 similarly includes an inner cowl 130 at a forward end 132. The outer cowl 126 and inner cowl 130 may assist in directing the flow of compressed air from the compressor section into or through one or more of the fuel air mixers 124.

Moreover, the inner and outer dome sections 116, 118 each include attachment portions configured to assist in mounting the combustor assembly 100 within the turbofan engine 10 (FIG. 1). For example, the outer dome section 118 includes an attachment extension 134 configured to be mounted to an outer combustor casing 136 and the inner dome section 116 includes a similar attachment extension 138 configured to attach to an annular support member 140 within the turbofan engine 10. In certain exemplary embodiments, the inner dome section 116 may be formed integrally as a single annular component, and similarly, the outer dome section 118 may also be formed integrally as a single annular component. It should be appreciated, however, that in other exemplary embodiments, the inner dome section 116 and/or the outer dome section 118 may alternatively be formed by one or more components being joined in any suitable manner. For example, with reference to the outer dome section 118, in certain exemplary embodiments, the outer cowl 126 may be formed separately from the outer dome section 118 and attached to the forward end 128 of the outer dome section 118 using, e.g., a welding process. Similarly, the attachment extension 134 may also be formed separately from the outer dome section 118 and attached to the forward end 128 of the outer dome section 118 using, e.g., a welding process. Additionally or alternatively, the inner dome section 116 may have a similar configuration.

With reference still to FIG. 2, the exemplary combustor assembly 100 further includes a heat shield 142 positioned around the fuel air mixer 124 depicted. For this embodiment, the exemplary heat shield 142 is attached to and extends between the outer dome section 118 and the inner dome section 116. The heat shield 142 is configured to protect certain components of the turbofan engine 10 (FIG. 1) from the relatively extreme temperatures of the combustion chamber 114.

For the embodiment depicted, the inner liner 102 and the outer liner 108 are each formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such liners 102, 108 may include silicon carbide, silicon, silica or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite and montmorillonite). CMC materials may have coefficients of thermal expansion in the range of about 1.3×10⁻⁶ in/in/° F. to about 3.5×10⁻⁶ in/in/° F. in a temperature of approximately 1000-1200° F.

By contrast, the annular dome, including the inner dome section 116 and outer dome section 118, may be formed of a metal, such as a nickel-based superalloy (having a coefficient of thermal expansion of about 8.3-8.5×10⁻⁶ in/in/° F. in a temperature of approximately 1000-1200° F.) or cobalt-based superalloy (having a coefficient of thermal expansion of about 7.8-8.1×10⁻⁶ in/in/° F. in a temperature of approximately 1000-1200° F.).

Referring still to FIG. 2, as noted above, the combustion gases 66 (FIG. 1) flow from the combustion chamber 114 into and through the turbine section of the turbofan engine 10 (FIG. 1) where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of turbine stator vanes and turbine rotor blades. A stage one (1) turbine nozzle 156 is depicted schematically in FIG. 2 positioned aft of the combustor assembly 100.

FIG. 3 provides a close up, schematic, cross-sectional view of an attachment point where the forward end 112 of the outer liner 108 is mounted to the outer dome section 118 within the slot 122 of the outer dome section 118. To allow for a relative thermal expansion between the outer liner 108 and the outer dome section 118, as well as between the inner liner 102 and the inner dome section 116 (FIG. 2), a plurality of mounting assemblies 144 are used to attach the outer liner 108 to the outer dome section 118 and the inner liner 102 to the inner dome section 116. More particularly, the mounting assemblies 144 attach the forward end 112 of the outer liner 108 to the outer dome section 118 within the slot 122 of the outer dome section 118 as shown in FIGS. 2 and 3 and the forward end 106 of the inner liner 102 to the inner dome section 116 within the slot 122 of the inner dome section 116 (FIG. 2). The slots 122 are defined by their respective domes. Moreover, the slots 122 receive the forward ends 106, 112 of the inner and outer liners 102, 108, respectively.

Referring particularly to the forward end 112 of the outer liner 108 and the outer dome section 118 depicted in FIG. 3, the outer dome section 118 includes a base plate 158 and a yolk 160. The base plate 158 and the yolk 160 are spaced along the radial direction R. The base plate 158 and the yolk 160 each extend substantially parallel to one another, which for the embodiment depicted is a direction substantially parallel to the axial direction A of the turbofan engine 10 (see also FIG. 2). Notably, the slot 122 is defined between the base plate 158 and the yolk 160. The slot 122 is further defined by the forward surface 121. Further, in certain exemplary embodiments, the yolk 160 may extend circumferentially with the outer dome section 118, tracking the base plate 158. With such a configuration, the slot 122 may be considered an annular slot. However, in other embodiments, the yolk 160 may include a plurality of circumferentially spaced tabs, each of the individual tabs of the yolk 160 defining individual segmented portions of the slot 122 with the base plate 158.

The exemplary mounting assembly 144 depicted includes the yolk 160 of the outer dome section 118 and the base plate 158 of the outer dome section 118. Moreover, the mounting assembly 144 includes a pin 162 and a bushing 164. The pin 162 includes a head 166 and a shank 168. The shank 168 extends through the yolk 160, the forward end 112 of the outer liner 108 (positioned in slot 122), and the base plate 158. A nut 170 is attached to a distal end of the shank 168 of the pin 162. In certain exemplary embodiments, the pin 162 may be configured as a bolt and the nut 170 may be rotatably engaged with a threaded portion of the pin 162 (at, e.g., the distal end of the shank 168) for tightening the mounting assembly 144. Alternatively, however, in other exemplary embodiments the pin 162 and nut 170 may have any other suitable configurations. In other exemplary embodiments, for instance, the pin 162 may include a shank 168 defining a substantially smooth cylindrical shape and the nut 170 may be configured as a clip.

Additionally, the bushing 164 is generally cylindrical in shape and is positioned around the shank 168 of the pin 162 within the slot 122. For the embodiment depicted, the bushing 164 is pressed between the yolk 160 and the base plate 158 by tightening the nut 170 on the pin 162. Moreover, for the embodiment depicted, the mounting assembly 144 includes a metal grommet 172 positioned around the bushing 164 and pin 162. The grommet 172 is positioned in a mounting opening 174 defined by the forward end 112 of the outer liner 108. The grommet 172 includes an outer collar 176 positioned adjacent to an outer surface 178 of the outer liner 108 and an inner collar 180 positioned adjacent to an inner surface 182 of the outer liner 108. The grommet 172 additionally includes a body 184. The metal grommet 172 may reduce an amount of wear on the forward end 112 of the outer liner 108 as the outer liner 108 moves inwardly and outwardly generally along the radial direction R relative to the outer dome section 118.

It should be appreciated, however, that although the forward end 112 of the outer liner 108 is attached to the outer dome section 118 using the exemplary mounting assembly 144 depicted and described herein, in other embodiments of the present disclosure, the mounting assembly 144 may have other suitable configurations, and further still in other embodiments, any other suitable attachment assembly may be used.

Referring still to FIG. 3, the forward end 112 of the outer liner 108 depicted further includes an axial interface surface 186 and a radial interface surface 188. The axial interface surface 186 is configured as a portion of the forward end 112 of the outer liner 108 facing the base plate 158 of the outer dome section 118, or more particularly, facing the inner surface 120 of the outer dome section 118. The radial interface surface 188 is configured as a portion of the forward end 112 of the outer liner 108 facing the forward surface 121 of the outer dome section 118. For the embodiment depicted, the axial interface surface 186 and inner surface 120 each extend in a direction parallel to the axial direction A, and the radial interface surface 188 and forward surface 121 each extend in a direction parallel to the radial direction R.

Moreover, as further shown in FIG. 3, the axial interface surface 186 defines a radial gap G_(R) with the inner surface 120 of the outer dome section 118 and the radial interface surface 188 defines an axial gap G_(A) with the forward surface 121 of the outer dome section 118. The combustor assembly 100 may be designed such that the radial and axial gaps G_(R), G_(A) allow for only a predetermined amount of airflow therethrough into the combustion chamber 114. Notably, allowing such a flow of air during operating conditions of the combustor assembly 100 may ensure relatively hot combustion gases within the combustion chamber 114 do not flow into and/or through the slot 122 of the outer dome section 118, potentially damaging certain components of the combustor assembly 100.

In addition to the airflow through the radial and axial gaps G_(R), G_(A), in some exemplary embodiments as will be explained more fully below, additional airflow can be provided to warm the forward end 112 of the outer liner 108 (as well as the forward end 106 of the inner liner 102 depicted in FIG. 2) to improve the thermal response (e.g., reduce the thermal lag) of the forward ends 112, 106 of the outer and inner liners 108, 102 during transient operation of the turbofan engine 10 (FIG. 1), and particularly during rapid power increases or bursts of the turbofan engine 10. In this way, the stress and strain on the outer and inner liners 108, 102 can be reduced during transient operation of the engine.

FIG. 4 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. Further, FIG. 4 depicts the forward end 112 of the outer liner 108 defining a plurality of blind warming openings 190. The diameter of the blind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm). The diameter of the mounting openings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). Thus, the mounting openings 174 have larger diameters than the blind warming openings 190. The blind warming openings 190 defined by the forward end 112 are operatively configured to increase the thermal response of the forward end 112 of the outer liner 108. As such, as will be explained more fully below, the stress and strain on the outer liner 108 can be reduced, thereby improving the durability of the outer liner 108. The forward end 106 of the inner liner 102 can likewise define blind warming openings 190, and thus, the durability of the inner liner 102 can likewise be improved.

As shown in FIG. 4, each of the blind warming openings 190 extend between an opening end 192 and a terminus end 194. The opening ends 192 of the blind warming openings 190 allow fluid (e.g., a warming airflow WA) to flow into and out of their respective openings and the terminus ends 194 are the dead ends or terminating points of the blind warming openings 190. For the depicted embodiment of FIG. 4, the blind warming openings 190 extend along the radial direction R from one of the outer surface 178 and the inner surface 182 and terminate prior to the opposing surface. In this way, the warming openings are blind openings (i.e., they do not extend through the forward end 112 of the outer liner 108). More particularly, as shown for the depicted embodiment of FIG. 4, one or more of the blind warming openings 190 extend along the radial direction R from the outer surface 178 and terminate prior to the opposing inner surface 182 and one or more of the blind warming openings 190 extend along the radial direction R from the inner surface 182 and terminate prior to the opposing outer surface 178. Thus, the blind warming openings 190 include outer surface openings 196 that extend along the radial direction R from the outer surface 178 and terminate prior to the inner surface 182 and the blind warming openings 190 include inner surface openings 198 that extend along the radial direction R from the inner surface 182 and terminate prior to the outer surface 178. The opening ends 192 of the outer surface openings 196 are defined by the outer surface 178 of the forward end 112 and the terminus ends 194 of the outer surface openings 196 are defined by the forward end 112 at some point along the radial thickness of the forward end 112. In contrast, the opening ends 192 of the inner surface openings 198 are defined by the inner surface 182 of the forward end 112 and the terminus ends 194 of the inner surface openings 198 are defined by the forward end 112 at some point along the radial thickness of the forward end 112.

As further shown in the circumferential cross section of FIG. 4, the forward end 112 defines a row of blind warming openings 190. More particularly, the row of blind warming openings 190 includes outer surface openings 196 that alternate with the inner surface openings 198 along the axial direction A. By alternating the outer surface openings 196 with the inner surface openings 198 along the axial direction A, the outer surface openings 196 and the inner surface openings 198 can be axially positioned near one another without compromising the structural integrity of the forward end 112 of the outer liner 108. In this way, efficient warming of the forward end 112 along its radial thickness and axial length can be achieved without, as noted previously, compromising the structural integrity of the forward end 112.

As further depicted in FIG. 4, a midline is defined midway between the outer and inner surfaces 178, 182 of the forward end 112 of the outer liner 108 along the radial direction R. As shown, for this embodiment, the blind warming openings 190 extend along the radial direction R from one of the outer and inner surfaces 178, 182 of the forward end 112 past the midline M. More particularly, as shown, the outer surface openings 196 extend from the outer surface 178 past the midline M and the inner surface openings 198 extend from the inner surface 182 past the midline M. More particularly still, the opening ends 192 of the outer surface openings 196 extend from the outer surface 178 and their terminus ends 194 extend past the midline M. Likewise, the opening ends 192 of the inner surface openings 198 extend from the inner surface 182 and their terminus ends 194 extend past the midline M. As the blind warming openings 190 extend along the radial direction R past the midline M in the depicted embodiment of FIG. 4, it is ensured that the forward end 112 is warmed along its radial thickness.

As further depicted in the exemplary embodiment of FIG. 4, the blind warming openings 190 are angled with respect to the radial direction R. For this embodiment, for instance, the opening ends 192 of the outer surface openings 196 and the inner surface openings 198 are positioned aft of their respective terminus ends 194 along the axial direction A. Stated differently, the opening ends 192 of the blind warming openings 190 are not aligned with their respective terminus ends 194 along the radial direction R. In this manner, the blind warming openings 190 are angled with respect to the radial direction R, and thus, the surface area of the sidewalls 199 defining the blind warming openings 190 is increased (e.g., as opposed to the sidewalls of blind warming openings oriented parallel with the radial direction R), which ultimately increases the rate of heat transfer or thermal response of the forward end 112.

Additionally, for the depicted embodiment of FIG. 4, the plurality of blind warming openings 190 are oriented at an angle θ with respect to the radial direction R. More particularly, for this embodiment, the outer surface openings 196 are angled at a negative forty-five degree (−45°) angle with respect to the radial direction R and the inner surface openings 198 are angled at a forty-five degree (45°) angle with respect to the radial direction R. Thus, the outer surface openings 196 are angled opposite the inner surface openings 198 with respect to the radial direction R. By angling the outer and inner surface openings 196, 198 opposite one another and alternating them along the axial direction A as noted above, the axial spacing of the blind warming openings 190 can be positioned so as to provide efficient warming of the forward end 112 along its radial thickness and axial length without compromising the structural integrity of the forward end 112. Moreover, as noted above, angling the blind warming openings 190 increases the surface area of the sidewalls 199 defining the openings, thereby increasing thermal response of the forward end 112. The blind warming openings 190 can be oriented in axial rows as shown in FIG. 4 along the circumferential direction C along the forward end 112 and between and/or about the mounting openings 174.

In other exemplary embodiments, the outer surface openings 196 are angled at a negative forty-five degree (−45°) angle with respect to the radial direction R and the inner surface openings 198 are likewise angled at a negative forty-five degree (−45°) angle with respect to the radial direction R. Thus, the outer surface openings 196 are not angled opposite the inner surface openings 198 with respect to the radial direction R. By angling the outer surface openings 196 at a negative angle with respect to the radial direction R, as warming airflow WA flows into the slot 122 between the yolk 160 and the forward end 112 of the outer liner 108 the warming airflow WA flows more directly into the outer surface openings 196. Thus, this may increase the rate of heat transfer of the forward end 112. Similarly, by angling the inner surface openings 198 at a negative angle with respect to the radial direction R, the warming airflow WA flowing into the slot 122 between the yolk 160 and the forward end 112, through the axial gap G_(A) (FIG. 3), and then aft between the inner surface 182 of the liner 108 and the base plate 158 flows more directly into the inner surface openings 198. This may increase the rate of heat transfer of the forward end 112.

In yet other exemplary embodiments, the warming openings 190 are angled at a sixty degree (60°) angle (positive or negative) with respect to the radial direction R, thereby further increasing the surface area of the sidewalls 199 defining the blind warming openings 190. In further embodiments, the blind warming openings 190 are angled at a seventy-five degree (75°) (positive or negative) angle with respect to the radial direction R. In some embodiments, the blind warming openings 190 are oriented at an angle that is greater than forty-five degrees (45°) (positive or negative) with respect to the radial direction R. It yet further embodiments, the blind warming openings 190 are oriented approximately along the radial direction R. By orienting the blind warming openings 190 along approximately along the radial direction R, the openings may be easier to machine into the forward end 112 of the outer liner 108 and may be less impactful to the structural integrity of the forward end 112.

During transient operation of the turbofan engine, and particularly during a rapid power increase or burst, the temperature of the generated combustion gases increases. The high temperature combustion gases scrub along the outer and inner liners, which rapidly heats the liners. However, as noted previously, the forward ends of the liners do not heat up as quickly. Stated differently, the forward ends of the liners thermally lag the other portions of the liners. Consequently, the liners may experience bending stress and strain due to the thermal lag of the forward ends. To reduce the thermal lag and thus the stress and strain on the liners, the blind warmings openings allow for a warming airflow to flow therein to warm the forward ends of the inner and outer liners. In this way, the stress and strain on the liners can be reduced, which as noted above, may improve the durability of the liners.

With reference to FIG. 4, a warming airflow WA may flow into the blind warming openings 190 as follows. As shown, during operation of turbofan engine 10 (FIG. 1), the warming airflow WA flows into the slot 122 defined by the outer dome 118. More particularly, the warming air WA flows between the yoke 160 and the outer surface 178 of the outer liner 108. In this example, the warming airflow WA is compressor discharge air (P3 air). As the warming airflow WA travels forward along the axial direction A deeper into the slot 122, some of the warming airflow WA flows into the blind warming openings 190, and more particularly the outer surface openings 196, as shown. The warming airflow WA flows into the outer surface openings 196 and convection heat transfer occurs. More specifically, the warming airflow WA flows into the outer surface openings 196 and exchanges heat with the sidewalls 199 of the openings, thus effectively warming the forward end 112 of the outer liner 108. Further, some of the warming airflow WA that enters slot 122 flows past the outer surface openings 196 and flows radially inward through the axial gap G_(A) (FIG. 3) between the forward surface 121 of the outer dome 118 and the radial interface surface 188 of the forward end 112. Thereafter, the warming airflow WA flows aft along the axial direction A between the inner surface 182 of the forward end 112 and the base plate 158 of the outer dome 118. Some of the warming airflow WA flows into the inner surface openings 198 and convection heat transfer occurs. That is, the warming airflow WA flows into the inner surface openings 198 and exchanges heat with the sidewalls 199 of the openings, thus effectively warming the forward end 112 of the outer liner 108. After flowing into and out of the inner surface openings 198, the warming airflow WA flows aft along the axial direction A through the radial gap G_(R) (FIG. 3) and into the combustion chamber 114. Flowing the warming airflow WA into and out of the blind warming openings 190, which in this embodiment includes the outer surface openings 196 and the inner surface openings 198, the forward end 112 of the outer liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of the forward end 112 is improved thereby reducing the bending stress and strain on the liner.

It will be appreciated that the forward end 106 of the inner liner 102 (FIG. 2) may be formed in the same or substantially the same manner as the forward end 112 of the outer liner 108 as described above. Further, it will be appreciated that the forward end 106 of the inner liner 102 may be attached to the inner dome section 116 in the same or substantially the same manner that the forward end 112 of the outer liner 108 is attached to the outer dome section 118.

FIG. 5 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. Further, FIG. 5 depicts the forward end 112 of outer liner 108 defining blind warming opening 190, which in the depicted embodiment of FIG. 5 is a blind warming slit 200. Like the blind warming openings depicted in the embodiment of FIG. 4, the warming slit 200 defined by the forward end 112 increases the thermal response of the forward end 112 of the outer liner 108. Although only one (1) warming slit 200 is shown in FIG. 5, in some embodiments the forward end 112 may define a plurality of warming slits 200. In particular, the forward end 112 of the outer liner 108 may define a plurality of warming slits 200 spaced apart from one another along the circumferential direction C. Moreover, although only one (1) slit is shown between the outer surface 178 and the inner surface 182 along the radial direction R, in some embodiments, the forward end 112 may define more than one warming slit 200 along the radial direction R. The forward end 106 of the inner liner 102 can likewise define blind warming slits 200. The diameter of the blind warming slits 200 may range between or about between 0.020 and 0.10 inches (0.508-2.54 mm). The diameter of the mounting openings 174 may range between or about between 0.400 and 0.800 inches (10.16-20.32 mm). As such, in such embodiments, the mounting openings 174 have larger diameters than the blind warming slits 200.

As shown in FIG. 5, the warming slit 200 extends between an opening end 202 and a terminus end 204 along the axial direction A. For this embodiment, the radial interface surface 188 defines the opening end 202 of the warming slit 200. Moreover for this embodiment, the warming slit 200 extends along the axial direction A aft of the mounting opening 174. In this way, during operation of turbofan engine 10 (FIG. 1), the forward end 112 is warmed by warming airflow WA from the radial interface surface 188 to a point along the axial length of the forward end 112 that is aft of the mounting assembly 144 that couples the forward end 112 of the outer liner 108 with the outer dome 118. As the area of the forward end 112 that is axially aligned with and forward of the mounting assemblies 144 are most susceptible to thermal lag, flowing warming airflow WA into the warming slit 200 along this length reduces the bending stress and strain on the liner during transient operation of the turbofan engine 10.

In some exemplary embodiments, the warming slit 200 is defined by the forward end 112 along the axial direction A from the radial interface surface 188 to a position that is aligned with or aft of an edge 143 along the axial direction A. The edge 143, as shown in FIG. 5, is a radial outermost point of a trailing edge 145 of heat shield 142. As the heat shield 142 generally facilitates the flow of combustion gases 66 downstream of the forward end 112 of the outer liner 108 (FIG. 2), the forward end 112 is susceptible to thermal lag from its radial interface surface 188 to a point along the axial direction A that is protected by the heat shield 142. Thus, by extending the warming slit 200 along this length effectively warms the forward end 112 during transient operation of the turbofan engine 10 (FIG. 1), and thus, bending stress and strain on the liner may be reduced.

In some exemplary embodiments, particularly where the combustor assembly 100 does not include heat shield 142 or similar structure, the warming slit 200 is defined by the forward end 112 along the axial direction A from the radial interface surface 188 to a position that is aligned with or aft of an edge 147 along the axial direction A. The edge 147, as shown in FIG. 5, is a radial outermost point of a trailing edge 149 of outer dome 118. As the outer dome 118 generally shields the forward end 112 of the outer liner 108 from combustion gases 66, the forward end 112 is susceptible to thermal lag from its radial interface surface 188 to a point along the axial direction A that is shielded by the outer dome 118. Thus, by extending the warming slit 200 along this length effectively warms the forward end 112 during transient operation of the turbofan engine 10 (FIG. 1), and thus, bending stress and strain on the liner may be reduced.

With reference to FIG. 5, warming airflow WA may flow into the blind warming openings 190, which in this embodiment is warming slit 200, as follows. As shown, during operation of turbofan engine 10 (FIG. 1), the warming airflow WA flows into the slot 122 defined by the outer dome 118. More particularly, the warming air WA flows between the yoke 160 and the outer surface 178 of the outer liner 108. In this example, the warming airflow WA is compressor discharge air (P3 air). After the warming airflow WA flows into the slot 122, the warming airflow WA flows around the most forward portion of the forward end 112 and flows into axial gap G_(A). As the warming airflow WA flows through the axial gap G_(A), some of the warming airflow WA flows into the opening end 202 of the warming slit 200. The warming airflow WA then travels forward along the axial direction A toward the terminus end 204 of the warming slit 200. As the warming airflow WA flows along the axial length of the warming slit 200, convection heat transfer occurs. That is, the warming airflow WA flowing through warming slit 200 exchanges heat with the walls 206 defining the warming slit 200, thus effectively warming the forward end 112 of the outer liner 108. After flowing into and out of the warming slit 200, the warming airflow WA flows aft along the axial direction A between the inner surface 182 of the outer liner 108 and the base plate 158 and through the radial gap G_(R) and into the combustion chamber 114. Flowing the warming airflow WA into and out of the blind warming openings 190, which in this embodiment is warming slit 200 or slits, the forward end 112 of the outer liner 108 is actively warmed. Accordingly, during transient bursts of the engine, the transient response rate of the forward end 112 may be improved, which may reduce the bending stress and strain on the liner.

In some embodiments, forward end 112 of outer liner 108 defines outer surface openings 196 (FIG. 4), inner surface openings 198 (FIG. 4), and warming slits 200 (FIG. 5). Moreover, it will be appreciated that the forward end 106 of the inner liner 102 can likewise include such configurations.

FIG. 6 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. FIG. 6 further depicts the outer dome 118 defining a plurality of impingement openings 210. In particular, the yolk 160 of the outer dome 118 defines the impingement openings 210. As shown, the yolk 160 extends between an outer surface 161 and an inner surface 163 along the radial direction R, and for this embodiment, the impingement openings 210 extend from the outer surface 161 to the inner surface 163 of the yolk 160. Thus, the impingement openings 210 are through holes that extend through the yolk 160.

As further shown in FIG. 6, the impingement openings 210 each extend between an outer end 212 and an inner end 214. The outer ends 212 of the impingement openings 210 are defined by the outer surface 161 of the yolk 160 and the inner ends 214 of the impingement openings 210 are defined by the inner surface 163 of the yolk 160. The outer end 212 of each impingement opening 210 has an outer diameter D_(O) and the inner end 214 of each impingement opening 210 has an inner diameter D_(I). For this embodiment, the outer diameters D_(O) of the impingement openings 210 are greater than the inner diameters D_(I). Stated differently, the impingement openings 210 taper as they extend radially inward from the outer surface 161 to the inner surface 163 of the yolk 160.

During operation of the turbofan engine 10 (FIG. 1), warming airflow WA (e.g., compressor discharge air) flows into the impingement openings 210. In particular, the warming airflow WA enters the impingement openings 210 through their respective outer ends 212. As the warming airflow WA flows inward along the radial direction R through the impingement openings 210, the pressure of the warming airflow increases due to the converging geometry of the impingement openings 210. Then, as the warming airflow WA exits the inner ends 214 of the impingement openings 210, the flow rate of the warming airflow WA rapidly increases thereby creating a high velocity or jet stream of warming airflow WA. The streams of warming airflow WA flow inward along the radial direction R and impinges on the outer surface 178 of the forward end 112 of the outer liner 108. By impinging air onto the forward end 112, the forward end 112 is actively warmed and thus the transient thermal response rate of the forward end 112 is increased. The increased transient thermal response of the forward end 112 may reduce the bending stress and strain on the liner.

In some alternative embodiments, the inner diameters D_(I) of the impingement openings 210 are greater than the outer diameters D_(O). Stated alternatively, the impingement openings 210 taper as they extend radially outward from the outer surface 161 to the inner surface 163 of the yolk 160. In such embodiments, the warming airflow WA flowing radially outward of the yolk 160 can more easily flow into the impingement openings 210, among other benefits. In yet other alternative embodiments, the inner diameters D_(I) of the impingement openings 210 are the same as the outer diameters D_(O). That is, the impingement openings 210 do not taper as they extend along the radial direction R. In further alternative embodiments, the impingement openings 210 are angled with respect to the radial direction R. For instance, as one example, the inner ends 214 of the impingement openings 210 may be positioned aft of their respective outer ends 212. In this way, the warming airflow WA may flow into the impingement openings 210 more easily.

FIG. 7 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. FIG. 7 further depicts the outer dome 118 defining impingement openings 210. Moreover, FIG. 7 also depicts an impingement jacket 220 attached to the outer dome 118. The impingement jacket 220 extends between a forward end 222 and an aft end (not shown) generally along the axial direction A. Generally, the impingement jacket 220 spans the axial length of the outer liner 108 (i.e., from the forward end 112 to the aft end 110 of the outer liner 108; see FIG. 2). As shown, the forward end 222 of the impingement jacket 220 is attached to the yolk 160 of the outer dome 118. The impingement jacket 220 may be attached to the yolk 160 in any suitable manner. The impingement jacket 220 may extend annularly about the circumferential direction C or may extend along the circumferential direction C in segments.

As further shown in FIG. 7, like the outer dome 118, the impingement jacket 220 defines a plurality of impingement openings 224. The impingement openings 224 may be configured in any suitable manner as noted above with respect to the impingement openings 210 defined by the yolk 160 of the outer dome 118. The diameter of the blind warming openings 190 may range between or about between 0.020 and 0.080 inches (0.508-2.032 mm). Accordingly, during operation of turbofan engine 10 (FIG. 1), warming airflow WA may flow through the impingement openings 224 defined by the impingement jacket 220 and impinge on the outer surface 178 of the forward end 112 of the outer liner 108.

FIG. 8 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. Moreover, FIG. 8 depicts metal grommet 172 of mounting assembly 144 positioned within mounting opening 174 and defining a plurality of warming passages 230. The warming passages 230 may be drilled into or otherwise machined into the grommet 172. As shown, the warming passages 230 defined by grommet 172 each extend between an outer end 232 and an inner end 234. For this embodiment, the warming passages 230 defined by grommet 172 extend along the radial direction R. The outer ends 232 of the warming passages 230 are defined by the outer collar 176 of the grommet 172, while the inner ends 234 of the warming passages 230 are defined by the inner collar 180 of the grommet 172. Further, as the warming passages 230 extend through the grommet 172, the warming passages 230 also extend through the body 184 of the grommet 172. Notably, as shown in FIG. 8, the warming passages 230 extend through the body 184 of the grommet 172 such that the warming passages 230 extend along a radial face 113 of the forward end 112 of the outer liner 108. In this way, during operation of the turbofan engine 10 (FIG. 1), warming airflow WA may flow through the warming passages 230 and warm the radial face 113 of the forward 112, thereby actively warming the forward end 112.

FIG. 9 provides a cross sectional view of the grommet 172 of FIG. 8 taken along line 9-9 of FIG. 8. As shown, the bushing 164 is generally cylindrical in shape and is positioned around the shank 168 of the pin 162. The bushing 164 and pin 162 are positioned within and extend through a grommet opening 173 defined by grommet 172. The grommet 172 is positioned in the mounting opening 174 defined by the forward end 112 of the outer liner 108. For this embodiment, as shown, the warming passages 230 are circumferentially spaced about the perimeter 185 of the body 184 of the grommet 172. Each warming passage 230 extends through the perimeter 185 of the body 184 as a cutout and has a semi-circle shaped radial cross section and is, as noted above, positioned or defined by grommet 172 such that the warming passages 230 are defined in part by the grommet 172 and defined in part by the forward end 112 of the outer liner 108. This, as noted above, may allow for warming airflow WA to pass therethrough and actively warm the radial face 113 and the forward end 112 more generally. Thus, by defining warming passages 230 in grommet 172, the transient thermal response rate of the forward end 112 can be increased, and accordingly, the bending stress and strain on the outer liner 108 may be reduced.

In some alternative embodiments, additionally or alternatively to manufacturing the grommet 172 with warming passages 230, the grommet 172 may be sized such that a gap is defined between the grommet 172 and the radial face 113 of the forward end 112. In this way, during operation of the turbofan engine 10 (FIG. 1), warming airflow WA may flow through the gap and warm the radial face 113, thereby actively warming the forward end 112 of the outer liner 108.

FIG. 10 provides a close up, cross-sectional view of one exemplary embodiment of combustor assembly 100 depicting forward end 112 of outer liner 108 attached to outer dome 118. Further, FIG. 10 depicts the forward end 112 having enhanced surfaces. In particular, for this embodiment, the outer surface 178 of the outer liner 108 has a heat transfer coefficient (HTC) enhanced surface 240 and the inner surface 182 of the outer liner 108 likewise has an HTC enhanced surface 240. By enhancing the surface of the outer surface 178 and inner surface 182 of the forward end 112 as shown in FIG. 10, the surface area of the outer surface 178 is increased and the surface area of the inner surface 182 is likewise increased. Additionally, the enhancement of the outer surface 178 and inner surface 182 causes more turbulent flow over or across the surfaces. Thus, by enhancing the surfaces of the forward end 112, the transient thermal response of the forward end 112 may be increased. Additionally, as the enhanced surfaces do not affect the chargeable flow through the turbofan engine 10 (FIG. 1), there is no penalty on engine efficiency or performance.

As one example, the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may be a rough bond coating. For instance, the coating may a suitable ceramic thermal and environmental barrier coating (TEBC) for CMC components. As another example, the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may have an undulating surface as shown in FIG. 10. As yet another example, the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may be a grit blasted coating. For instance, the outer and/or inner surfaces 178, 182 of the forward end 112 may be blasted with a coarse grit. As a further example, the enhanced surface 240 of the outer surface 178 and/or the inner surface 182 may have a plurality of bumps having a height of about ten (10) millimeters.

Although the exemplary embodiments of the present disclosure were mainly discussed and illustrated using the outer liner and outer dome section of the combustor assembly, it will be appreciated that each exemplary aspect disclosed herein is applicable to the inner liner and inner dome section of the combustor assembly.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims. 

What is claimed is:
 1. A combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction, the combustor assembly comprising: a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, wherein the forward end of the liner defines a plurality of mounting openings spaced along the circumferential direction, and wherein the forward end of the liner defines one or more blind warming openings.
 2. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein each of the plurality of blind warming openings extend along the radial direction from one of the outer surface and the inner surface and terminate prior to the opposing surface.
 3. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein one or more of the plurality of blind warming openings extend along the radial direction from the outer surface and terminate prior to the inner surface and one or more of the plurality of blind warming openings extend along the radial direction from the inner surface and terminate prior to the outer surface.
 4. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein a midline is defined midway between the outer and inner surfaces along the radial direction, wherein the plurality of blind warming openings extend along the radial direction from one of the inner and outer surfaces of the forward end past the midline.
 5. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein the plurality of blind openings comprise a plurality of outer surface openings that each extend along the radial direction from the outer surface and terminate prior to the inner surface and a plurality of inner surface openings that each extend along the radial direction from the inner surface and terminate prior to the outer surface, and wherein the outer surface openings alternate with the inner surface openings along the axial direction.
 6. The combustor assembly of claim 5, wherein the plurality of outer surface openings extend between an opening end and a terminus end and the plurality of inner surface openings extend between an opening end and a terminus end, and wherein the opening ends of the outer surface openings and the inner surface openings are positioned aft of their respective terminus ends along the axial direction.
 7. The combustor assembly of claim 1, wherein the plurality of blind warming openings are angled with respect to the radial direction.
 8. The combustor assembly of claim 1, wherein the plurality of blind warming openings comprises one or more warming slits that extend between an opening end and a terminus end along the axial direction.
 9. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, the forward end further having a radial interface surface extending between and connecting the outer surface with the inner surface, and wherein the plurality of blind warming openings comprises one or more warming slits that extend along the axial direction between an opening end defined by the radial interface surface of the forward end and a terminus end.
 10. The combustor assembly of claim 9, wherein the one or more warming slits extend along the axial direction aft of the plurality of the mounting openings.
 11. The combustor assembly of claim 1, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein at least one of the outer surface and the inner surface comprise a surface enhancement, wherein the surface enhancement is at least one of a rough bond coating, a plurality of bumps, and a grit blasted coating.
 12. The combustor assembly of claim 1, wherein the liner is formed of a ceramic matrix composite (CMC) material.
 13. A combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction, the combustor assembly comprising: a dome comprising a yolk and a base plate spaced outward from the yolk along the radial direction, the yolk and the base plate defining a slot; and a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot; wherein the yolk defines a plurality of impingement openings extending therethrough such that a warming airflow impinges on the forward end of the liner during operation of the gas turbine engine.
 14. The combustor assembly of claim 13, further comprising: an impingement jacket extending between a forward end and an aft end, the forward end of the impingement jacket attached to the yolk of the dome, wherein the forward end of the impingement jacket defines a plurality of impingement openings such that a warming airflow impinges on the forward end of the liner during operation of the gas turbine engine.
 15. The combustor assembly of claim 13, wherein the impingement openings each extend between an outer end and an inner end, wherein the outer end has a diameter that is greater than a diameter of the inner end of one or more of the plurality of impingement openings.
 16. A combustor assembly for a gas turbine engine defining an axial direction, a radial direction, and a circumferential direction, the combustor assembly comprising: a dome defining a slot; a liner at least partially defining a combustion chamber and extending between an aft end and a forward end, the forward end of the liner received within the slot of the dome, the forward end of the liner defining a plurality of mounting openings spaced along the circumferential direction; and a plurality of mounting assemblies for coupling the forward end of the liner with the dome, each of the plurality of mounting assemblies comprising: a pin extending through one of the plurality of mounting openings; and a grommet positioned within one of the plurality of mounting openings and defining a grommet opening, the pin received within the grommet opening, and wherein the grommet defines one or more warming passages.
 17. The combustor assembly of claim 16, wherein the forward end of the liner extends between an outer surface and an opposing inner surface along the radial direction, and wherein the grommet comprises an outer collar positioned adjacent the outer surface of the liner and an inner collar positioned adjacent the inner surface of the liner, and wherein the plurality of warming passages extend through the grommet between the outer collar and the inner collar of the grommet.
 18. The combustor assembly of claim 16, wherein the grommet defines the plurality of warming passages such that the warming passages are defined in part by the grommet and defined in part by the forward end of the outer liner.
 19. The combustor assembly of claim 16, wherein the grommet comprises an outer collar, and inner collar, and a body extending therebetween along the radial direction, the body having a perimeter, and wherein the plurality of warming passages extend through the grommet from the outer collar to the inner collar along the radial direction and define cutouts along the perimeter of the body of the grommet.
 20. The combustor assembly of claim 16, wherein the grommet is sized such that a warming airflow can flow between the grommet and the liner during operation of the gas turbine engine. 